This disclosure relates to a gas turbine engine airfoil. More particularly, the disclosure relates to a cooling arrangement within a cooling passage of the airfoil.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades.
Many blades and vanes, blade outer air seals, turbine platforms, and other components include internal cooling passages having turns that provide a serpentine shape, which create undesired pressure losses. Some of the cooling passages may include portions having turbulence promoters that enhance the cooling effects of the cooling flow through the cooling passage, in particular, at the trailing edge portion of the airfoil. Typical designs carry either a full rib of cross-overs or pedestals extending radially within the trailing edge portion.